Airfoil cooling using non-line of sight holes

ABSTRACT

An airfoil for a gas turbine engine is provided that includes a first portion formed from a first plurality of plies of a ceramic matrix composite material and defining an inner surface of the airfoil, as well as a second portion formed from a second plurality of plies of a ceramic matrix composite material and defining an outer surface of the airfoil. The first portion and the second portion define a non-line of sight cooling aperture extending from the inner surface to the outer surface of the airfoil. In one embodiment, a surface angle that is less than 45° is defined between a second aperture and the outer surface. A method for forming an airfoil for a gas turbine engine also is provided.

FIELD OF THE INVENTION

The present subject matter relates generally to a gas turbine engine, ormore particularly to features for cooling internal components of gasturbine engines. Most particularly, the present subject matter relatesto non-line of sight cooling holes for gas turbine engine airfoils.

BACKGROUND OF THE INVENTION

A gas turbine engine generally includes a fan and a core arranged inflow communication with one another. Additionally, the core of the gasturbine engine generally includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. In operation, air is provided from the fan to an inlet of thecompressor section where one or more axial compressors progressivelycompress the air until it reaches the combustion section. Fuel is mixedwith the compressed air and burned within the combustion section toprovide combustion gases. The combustion gases are routed from thecombustion section to the turbine section. The flow of combustion gasesthrough the turbine section drives the turbine section and is thenrouted through the exhaust section, e.g., to atmosphere.

In general, turbine performance and efficiency may be improved byincreased combustion gas temperatures. However, increased combustiontemperatures can negatively impact the gas turbine engine components,for example, by increasing the likelihood of material failures. Thus,while increased combustion temperatures can be beneficial to turbineperformance, some components of the gas turbine engine may requirecooling features or reduced exposure to the combustion gases to decreasethe negative impacts of the increased temperatures on the components.

Film cooling gas turbine engine components, e.g., by directing a flow ofcooler fluid over the surface of the component, can help reduce thenegative impacts of elevated combustion temperatures. For example,cooling apertures may be provided throughout a component; the coolingapertures may allow a flow of cooling fluid from within the component tobe directed over the outer surface of the component. However, in areasof high curvature of the component, it can be difficult to direct theflow of cooling fluid from the cooling apertures over the outer surfaceof the component to form a cooling film of fluid. Further, known methodsof forming cooling apertures, e.g., by boring or otherwise machiningapertures in the component, can be ineffective in producing optimalcooling aperture lengths for controlling bore cooling and in producingcooling apertures having optimal fluid exit surface angles. In addition,known methods of machining cooling apertures are prone to through-holescarfing and often present challenges to properly positioning thecooling apertures.

Therefore, improved cooling features for gas turbine components thatovercome one or more disadvantages of existing cooling features would bedesirable. In particular, an airfoil for a gas turbine engine havingfeatures for reducing an angle between a cooling aperture and an outersurface of the airfoil to reduce a surface angle of cooling fluidexiting the cooling aperture would be beneficial. Further, an airfoilhaving a cooling aperture including a change in direction between afirst section and a second section of the cooling aperture would beadvantageous. Additionally, a method for forming an airfoil for a gasturbine engine, the airfoil having features for improved surface coolingof the airfoil, would be useful.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, an airfoil for agas turbine engine is provided. The airfoil includes a first portiondefining an inner surface of the airfoil. The first portion is formedfrom a first plurality of plies of a ceramic matrix composite material.The airfoil also includes a second portion defining an outer surface ofthe airfoil. The second portion is formed from a second plurality ofplies of a ceramic matrix composite material. The first portion and thesecond portion define a non-line of sight cooling aperture extendingfrom the inner surface to the outer surface of the airfoil.

In another exemplary embodiment of the present disclosure, an airfoilfor a gas turbine engine is provided. The airfoil includes a firstportion defining an inner surface of the airfoil. The first portion isformed from a first plurality of plies of a ceramic matrix compositematerial. The airfoil also includes a second portion defining an outersurface of the airfoil. The second portion is formed from a secondplurality of plies of a ceramic matrix composite material. The firstportion defines a first aperture and the second portion defines a secondaperture. The first aperture and the second aperture extend alongdifferent directions and define a cooling aperture forming a continuouspathway through the airfoil from the inner surface to the outer surface.Further, a surface angle is defined between the second aperture and theouter surface. The surface angle is less than 45°.

In a further exemplary embodiment of the present disclosure, a methodfor forming an airfoil for a gas turbine engine is provided. The methodincludes laying up a first plurality of plies of a ceramic matrixcomposite material; processing the first plurality of plies to form afirst portion of the airfoil; laying up a second plurality of plies of aceramic matrix composite material, the second plurality of plies laid upon the first portion of the airfoil; and processing the second pluralityof plies and the first portion to form a second portion of the airfoiladjacent the first portion. The first and second portions define anon-line of sight cooling aperture.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 provides a schematic cross-sectional view of an exemplary gasturbine engine according to various embodiments of the present subjectmatter.

FIG. 2 provides a side, perspective view of a turbine rotor bladeaccording to an exemplary embodiment of the present subject matter.

FIG. 3 provides a perspective view of a turbine nozzle segment accordingto an exemplary embodiment of the present subject matter.

FIG. 4 provides a cross-section view of a portion of an airfoil of theturbine nozzle segment, taken along the line 4-4 of FIG. 3, according toan exemplary embodiment of the present subject matter.

FIG. 5 provides a cross-section view of the portion of the airfoil ofthe turbine nozzle segment of FIG. 4, taken along the line 5-5 of FIG. 4through a cooling aperture of the airfoil, according to an exemplaryembodiment of the present subject matter.

FIG. 6 provides a plan view of the cooling aperture of FIG. 5 accordingto an exemplary embodiment of the present subject matter.

FIG. 7 provides the cross-section view through the cooling aperture ofthe airfoil shown in FIG. 5 according to another exemplary embodiment ofthe present subject matter.

FIG. 8 provides a plan view of the cooling aperture of FIG. 7 accordingto an exemplary embodiment of the present subject matter.

FIG. 9 provides a cross-section view of a portion of the airfoil of theturbine nozzle segment of FIG. 3 according to an exemplary embodiment ofthe present subject matter.

FIG. 10 provides a cross-section view of a portion of the airfoil of theturbine nozzle segment of FIG. 3 according to another exemplaryembodiment of the present subject matter.

FIG. 11 provides a chart illustrating a method for forming an airfoil ofa gas turbine engine according to an exemplary embodiment of the presentsubject matter.

DETAILED DESCRIPTION OF THE INVENTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first,” “second,” and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows and “downstream” refers to thedirection to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1, the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1, the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference) and a radial direction R. In general, the turbofan 10includes a fan section 14 and a core turbine engine 16 disposeddownstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22.

For the depicted embodiment, fan section 14 includes a variable pitchfan 38 having a plurality of fan blades 40 coupled to a disk 42 in aspaced apart manner. As depicted, fan blades 40 extend outward from disk42 generally along the radial direction R. Each fan blade 40 isrotatable relative to disk 42 about a pitch axis P by virtue of the fanblades 40 being operatively coupled to a suitable actuation member 44configured to vary the pitch of the fan blades 40. Fan blades 40, disk42, and actuation member 44 are together rotatable about thelongitudinal axis 12 by LP shaft 36 across a power gear box 46. Thepower gear box 46 includes a plurality of gears for stepping down therotational speed of the LP shaft 36 to a more efficient rotational fanspeed.

Referring still to the exemplary embodiment of FIG. 1, disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated thatnacelle 50 may be configured to be supported relative to the coreturbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50 mayextend over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersturbofan 10 through an associated inlet 60 of the nacelle 50 and/or fansection 14. As the volume of air 58 passes across fan blades 40, a firstportion of the air 58 as indicated by arrows 62 is directed or routedinto the bypass airflow passage 56 and a second portion of the air 58 asindicated by arrows 64 is directed or routed into the LP compressor 22.The ratio between the first portion of air 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the secondportion of air 64 is then increased as it is routed through the highpressure (HP) compressor 24 and into the combustion section 26, where itis mixed with fuel and burned to provide combustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

In some embodiments, components of turbofan engine 10, particularlycomponents within hot gas path 78, may comprise a ceramic matrixcomposite (CMC) material, which is a non-metallic material having hightemperature capability. Exemplary CMC materials utilized for suchcomponents may include silicon carbide, silicon, silica, or aluminamatrix materials and combinations thereof. Ceramic fibers may beembedded within the matrix, such as oxidation stable reinforcing fibersincluding monofilaments like sapphire and silicon carbide (e.g.,Textron's SCS-6), as well as rovings and yarn including silicon carbide(e.g., Nippon Carbon's NICALON®, Ube Industries' TYRANNO®, and DowCorning's SYLRAMIC®), alumina silicates (e.g., Nextel's 440 and 480),and chopped whiskers and fibers (e.g., Nextel's 440 and SAFFIL®), andoptionally ceramic particles (e.g., oxides of Si, Al, Zr, Y, andcombinations thereof) and inorganic fillers (e.g., pyrophyllite,wollastonite, mica, talc, kyanite, and montmorillonite). As furtherexamples, the CMC materials may also include silicon carbide (SiC) orcarbon fiber cloth.

CMC materials may be used for various components of the engine, forexample, airfoils in the turbine, compressor, and/or fan regions. Thecompressor and turbine generally include rows of airfoils that arestacked axially in stages. Each stage includes a row ofcircumferentially spaced stator vanes and a rotor assembly that rotatesabout centerline 12 of engine 10. Turbine nozzles, comprising statorvanes extending between inner and outer bands, direct the hot combustiongas in a manner to maximize extraction at the adjacent downstreamturbine blades. In various embodiments of engine 10, the nozzles and/orturbine blades, including their associated airfoils, may be CMCcomponents. Of course, other components of turbine engine 10 also may beformed from CMC materials.

Referring now to FIG. 2, a side, perspective view of a turbine rotorblade 74 is provided according to an exemplary embodiment of the presentsubject matter. As previously described, LP turbine 30 includessequential stages of turbine stator vanes 72 coupled to outer casing 18and turbine rotor blades 74 coupled to shaft or spool 36. Each blade 74includes an airfoil 80 having a pressure side 82 opposite a suction side84. Opposite pressure and suction sides 82, 84 of each airfoil 80 extendradially along a blade span S from a blade root 86 to a blade tip 87. Asdepicted, blade root 86 is the radially innermost portion of blade 74and the blade tip is the radially outermost portion of blade 74.Moreover, as further shown in FIG. 2, pressure and suction sides 82, 84of airfoil 80 extend axially between a leading edge 88 and an oppositetrailing edge 90. Airfoil 80 defines a chord C extending axially betweenopposite leading and trailing edges 88, 90. Moreover, airfoil 80 definesa width W between pressure side 82 and suction side 84. The width W ofairfoil 80 may vary along the span S.

Each blade 74 is coupled to shaft or spool 36 via blade root 86. Moreparticularly, blade root 86 is coupled to a turbine rotor disk (notshown), which in turn is coupled to shaft or spool 36 (FIG. 1). It willbe readily understood that, as is depicted in FIG. 2 and is generallywell-known in the art, blade root 86 may define a projection 89 having adovetail or other shape for receipt in a complementarily shaped slot inthe turbine rotor disk to couple blade 74 to the disk. Of course, eachblade 74 may be coupled to the turbine rotor disk and/or shaft or spool36 in other ways as well. In any event, blades 74 are coupled to theturbine rotor disks such that a row of circumferentially adjacent blades74 extends radially outward from the perimeter of each disk, i.e.,adjacent blades 74 within a blade row are spaced apart from one anotheralong a circumferential direction M and each blade 74 extends from thedisk along the radial direction R. As such, the turbine rotor disk andouter casing 18 form an inner end wall and an outer end wall,respectively, of hot gas path 78 through the turbine assembly.

Referring now to FIG. 3, a perspective view is provided of a turbinenozzle segment. A turbine stator is formed by a plurality of turbinenozzle segments that are abutted at circumferential ends to form acomplete ring about centerline 12. Each nozzle segment may comprise oneor more vanes, such as vanes 68 of HP turbine 28 or vanes 72 of LPturbine 30, that extend between an outer band and an inner band aspreviously described. FIG. 3 depicts an exemplary turbine nozzle segment67 of HP turbine 28. Nozzle segment 67 includes outer band 67 a andinner band 67 b, between which extends stator vanes 68. Each stator vane68 includes an airfoil 80, which has the same features as airfoil 80described above with respect to blade 74. For example, airfoil 80 ofvane 68 has a pressure side 82 opposite a suction side 84. Oppositepressure and suction sides 82, 84 of each airfoil 80 extend radiallyalong a span from a vane root at inner band 67 b to a vane tip at outerband 67 a. Moreover, pressure and suction sides 82, 84 of airfoil 80extend axially between a leading edge 88 and an opposite trailing edge90. Airfoil 80 further defines a chord extending axially betweenopposite leading and trailing edges 88, 90. Moreover, airfoil 80 definesa width between pressure side 82 and suction side 84, which may varyalong the span.

It will be appreciated that, although airfoil 80 of vane 68 may have thesame features as airfoil 80 of blade 74, airfoil 80 of vane 68 may havea different configuration than airfoil 80 of blade 74. As an example,the span of airfoil 80 of vane 68 may be larger or smaller than the spanof airfoil 80 of blade 74. As another example, the width and/or chord ofairfoil 80 of vane 68 may differ from the width and/or chord of airfoil80 of blade 74. Additionally or alternatively, airfoils 80 of LP statorvanes 72 and/or airfoils 80 of HP turbine rotor blades 70 may differ insize, shape, and/or configuration from airfoils 80 of HP stator vanes 68and LP turbine rotor blades 74. However, it also should be understoodthat, while airfoils 80 may differ in size, shape, and/or configuration,the subject matter described herein may be applied to any airfoil withinengine 10, as well as other suitable components of engine 10.

FIG. 4 provides a cross-sectional view of a portion of airfoil 80 ofstator vane 68, taken along the line 4-4 of FIG. 3, according to anexemplary embodiment of the present subject matter. FIG. 5 provides across-sectional view of a portion of airfoil 80, taken along the line5-5 of FIG. 4, according to an exemplary embodiment of the presentsubject matter. As illustrated, airfoil 80 comprises a first portion 92and a second portion 94. First portion 92 is fabricated from a firstplurality of plies 96 of a CMC material and second portion 94 isfabricated from a second plurality of plies 98 of a CMC material. Firstportion 92 defines an inner surface 100 of airfoil 80, and secondportion 94 defines an outer surface 102 of airfoil 80.

Referring still to FIGS. 4 and 5, airfoil 80 defines cooling apertures120 for providing a flow of cooling fluid over outer surface 102 ofairfoil 80. Each cooling aperture 120 comprises a first section and asecond section. More particularly, first portion 92 of airfoil 80defines a first aperture 104 therethrough; first aperture 104 is thefirst section of cooling aperture 120. First aperture 104 has a firstend 106 defined at inner surface 100 and a second end 108 defined atsecond portion 94 of airfoil 80. First and second ends 106, 108 arespaced apart by a first length L₁ (FIG. 5). Further, first aperture 104extends from its first end 106 to its second end 108 along a firstdirection D₁.

In addition, second portion 94 of airfoil 80 defines a second aperture110 therethrough; second aperture 110 is the second section of coolingaperture 120. Second aperture 110 has a first end 112 defined adjacentsecond end 108 of first aperture 104. Second aperture 110 also has asecond end 114 defined at outer surface 102 of airfoil 80. First andsecond ends 112, 114 are spaced apart by a second length L₂ (FIG. 6).Second aperture 110 extends from its first end 112 to its second end 114along a second direction D₂.

Referring back to FIG. 4, first portion 92 defines a cavity 116 thatreceives a flow of cooling fluid F, e.g., a flow of pressurized airdiverted from HP compressor 24. The fluid flow F received within cavity116 generally is cooler than the combustion gases flowing against orover outer surface 102 of airfoil 80. Each cooling aperture 120,extending from cavity 116 to outer surface 102 via first aperture 104and second aperture 110, forms a continuous pathway in fluidcommunication with cavity 116 to facilitate flowing cooling fluid F fromcavity 116 to outer surface 102. As such, the flow of cooling fluid Fover outer surface 102 can help reduce the temperatures to which outersurface 102 is exposed.

As shown in FIG. 4, and more clearly in FIG. 5, first direction D₁ isdifferent from second direction D₂. The change in direction betweenfirst aperture 104 and second aperture 110 provides a non-line of sightcooling aperture 120, i.e., cavity 116 of airfoil 80 cannot be viewedfrom the exterior of the airfoil via cooling aperture 120, which extendsfrom inner surface 100 to outer surface 102 of airfoil 80. The change indirection allows an angle α between second aperture 110 and outersurface 102 of airfoil 80 to be reduced such that a reduced surfaceangle α is defined between second aperture 110 and outer surface 102.More particularly, a single direction cooling aperture requires acertain minimum surface angle α for the cooling aperture or hole to beable to extend from inner surface 100 to outer surface 102 and therebypermit a flow of cooling fluid from cavity 116 to outer surface 102 ofairfoil 80. That is, in single direction cooling hole configurations, ifthe angle between the cooling hole and the outer surface of the airfoilis below the minimum value (i.e., too small), the cooling hole will notbe able to extend from the outer surface to the inner surface to accessthe cooling fluid. However, by utilizing first aperture 104 to accesscavity 116 and second aperture 110 to access outer surface 102, wherefirst aperture 104 and second aperture 110 extend along differentdirections yet form a continuous pathway from cavity 116 to outersurface 102, the cooling aperture 120 of the present subject matterpermits small surface angles a between the cooling aperture and theouter surface of the airfoil. Preferably, the surface angle α is nogreater than, or less than, 45°; that is, the surface angle α definedbetween second aperture 110 and outer surface 102 is between 0° and 45°.A reduced surface angle can better direct the flow of cooling fluid Fexiting cooling aperture 120 along outer surface 102 and thereby improvefilm effectiveness and, thus, surface cooling of airfoils of gas turbineengine 10.

Further, it will be readily understood that any number of coolingapertures 120 comprising first and second apertures 104, 110 may be usedthroughout airfoil 80. More specifically, as represented by the dashedlines in FIG. 4, cooling apertures 120 may be defined at variouslocations along the span S of airfoil 80, i.e., spaced apart overairfoil 80 generally along the radial direction R. Moreover, the changein direction between first aperture 104 and second aperture 110 allowscooling apertures 120 to be particularly beneficial for areas of airfoil80 having a high curvature, such as leading edge 88, where it may bedifficult to direct a flow of cooling fluid over outer surface 102.However, although cooling apertures 120 may be particularly beneficialalong leading edge 88 of airfoil 80, cooling apertures 120 may besuitable for any location on airfoil 80 and, for example, may be definedalong trailing edge 90 as well as over pressure and suction sides 82, 84of airfoil 80. Further, cooling apertures 120 also may be defined on arotor blade platform, as well as along an inner and/or outer wall orband of a nozzle, such as along inner band 67 b and/or outer band 67 aof nozzle 67 shown in FIG. 3. As such, cooling apertures 120 may bespaced apart generally along the axial direction A as well as the radialdirection R. Further, the use of multiple cooling apertures 120 atmultiple locations of airfoil 80 may help enhance the surface coolingprovided by the cooling fluid flowing from each aperture 120.

Additionally, it will be appreciated that, although described herein ashaving two sections (first aperture 104 being the first section andsecond aperture 110 being the second section), in other embodimentscooling apertures 120 may have more than two sections. For example,airfoil 80 may comprise three portions—a first portion defining innersurface 100 and cavity 116, a second portion defined outside the firstportion (i.e., defined adjacent the first portion outward of the innersurface 100), and a third portion defined outside the second portion(i.e., defined adjacent the second portion outward of the first portion)and defining outer surface 102. Each portion of airfoil 80 may define anaperture, and each aperture may extend along a different direction or,at the least, extend along two different directions. The aperture of oneportion may be defined adjacent the aperture of the adjacent portion toform a continuous pathway, and, thus, a cooling aperture 120, fromcavity 116 to outer surface 102. Because the apertures extend along atleast two different directions, the cooling aperture 120 is a non-lineof sight aperture, with a change in direction of the aperture 120between inner surface 100 of airfoil 80 and outer surface 102.

Moreover, as will be appreciated from the illustrations of FIGS. 4 and5, cooling apertures 120 may vary in length. That is, one coolingaperture 120 may have a different overall length, e.g., the sum of firstlength L₁ and second length L₂, than another cooling aperture 120. Inother embodiments, the length L₁ of first aperture 104 and/or the lengthL₂ of second aperture 110 may vary among cooling apertures 120. Byvarying the length of cooling apertures 120, the cooling can becontrolled, i.e., the length of cooling apertures 120 is customizable toaccommodate different cooling requirements for different locations ofairfoil 80 and to control cooling local to cooling apertures 120.

As illustrated in FIGS. 5 through 8, each ply of second plurality ofplies 98 may be cut to define a cut-out 118. The plurality of cut-outs118 define second aperture 110. Further, although illustrated as amachined bore through first portion 92, it will be appreciated thatfirst aperture 104 similarly may be defined by cut-outs in firstplurality of plies 96. That is, as described with respect to secondaperture 110, each ply of first plurality of plies 96 may be cut todefine a cut-out 118 in the ply, and the plurality of cut-outs 118 infirst plies 96 may define first aperture 104. First and second apertures104, 110 may be defined in other ways as well.

Referring particularly to FIG. 6, a plan view is provided of coolingaperture 120 according to an exemplary embodiment of the present subjectmatter. As depicted in FIG. 6, first aperture 104 has a firstcross-sectional area A₁ that is constant over its length L₁ from innersurface 100 to second portion 94. Further, in the depicted embodiment,cut-outs 118 forming second aperture 110 have a second cross-sectionalarea A₂ that increases along the length L₂ of second aperture 110, i.e.,along second direction D₂ from first end 112 at first portion 92 tosecond end 114 at outer surface 102. Conversely, the cross-sectionalarea A₂ of second aperture 110 decreases from its second end 114 to itsfirst end 112. In exemplary embodiments, such as the embodiment shown inFIG. 5, the cross-sectional area A₂ of second aperture 110 may increasealong second direction D₂ from about two times greater or larger thanthe cross-sectional area A₁ of first aperture 104 to about twelve timesgreater or larger than area A₁. That is, the cross-sectional area A₂ ofthe cut-out 118 at outer surface 102 of airfoil 80 may be approximatelytwelve times greater or larger than the cross-sectional area A₁ of firstaperture 104 and the cut-out 118 at first end 112 of second aperture 110may be about twice the area A₁ of first aperture 104. As such, firstaperture 104 has a minimum cross-sectional area A₁ and second aperture110 has a minimum cross-sectional area A₂. Further, in such embodiments,the exit area A₂ of cut-out 118 at outer surface 102 may be greater thanthe minimum cross-sectional area A₁, e.g., larger than the smallestcross-sectional area value of first aperture 104 as just described. Inother embodiments, first aperture 104 may have a varying cross-sectionalarea A₁ and second aperture 110 may have a constant cross-sectional areaA₂. In such embodiments, the minimum cross-sectional area A₁ of firstaperture 104 also may be smaller than the minimum cross-sectional areaA₂ of second aperture 110.

Although with the top, plan view of FIG. 6 illustrates centers of firstaperture 104 and second aperture 110 aligned with respect to an axisH-H, first aperture 104 and second aperture 110 need not be aligned. Forexample, as shown in FIGS. 7 and 8, first and second apertures 104, 110may be defined such that a centerline CL₂ of second aperture 110 isaligned with axis H-H but a centerline CL₁ of first aperture 104 isaligned with a different axis I-I.

Moreover, each cut-out 118 may have a center 122, and in embodimentssuch as shown in FIGS. 5 and 6, each cut-out 118 in plies 98 may bealigned such that the center 122 of each cut-out 118 is aligned witheach center 122 along second direction D₂. In other embodiments,however, cut-outs 118 may be offset from one another with respect tosecond direction D₂ and/or with respect to axis H-H. For example, asillustrated in FIGS. 9 and 10, a center 122 of cut-out 118 of one plymay be offset from the center 122 of cut-out 118 of the successiveand/or preceding ply. That is, centers 122 may not be aligned as shownin FIGS. 4 and 5 but, rather, may be misaligned as shown in FIGS. 9 and10.

As further illustrated in FIGS. 9 and 10, cut-outs 118 of adjacent pliesof second plurality of plies 98 may have different shapes and/orcross-sections. More particularly, FIG. 5 illustrates cut-outs 118 ashaving the same general rectangular or pill shape and uniformlyincreasing or decreasing in cross-sectional area. However, in otherembodiments, adjacent cut-outs 118 may have the same cross-sectionalarea and/or different shapes. In still other embodiments, any change incross-sectional area of cut-outs 118 may be non-uniform. As an example,a cut-out 118 in ply A may have a cross-sectional area A_(A), a cut-out118 in adjacent ply B may have a cross-sectional area A_(B), and acut-out 118 in ply C (adjacent ply B) may have a cross-sectional areaA_(C). Referring to FIG. 9, cross-sectional area A_(A) of cut-out 118 inply A may be approximately equal to the cross-sectional area A_(B) ofcut-out 118 in ply B, but cross-sectional area A_(C) of cut-out 118 inply C may be smaller than cross-sectional areas A_(A) and A_(B). Asanother example shown in FIG. 10, cross-sectional area A_(B) of cut-out118 in ply B may be larger or greater than both cross-sectional areaA_(A) of cut-out 118 in ply A and cross-sectional area A_(C) of cut-out118 in ply C. Thus, in the embodiments shown in FIGS. 9 and 10, thechange in cross-sectional area of cut-outs 118 from ply A to ply C alongsecond direction D₂ is non-uniform.

As previously described, airfoil 80 may be formed from first pluralityof plies 96 and second plurality of plies 98. FIG. 11 provides a chartillustrating an exemplary method 800 for fabricating airfoil 80. At step802 of exemplary method 800, first plurality of plies 96 are laid up,e.g., in or on a layup tool, mandrel, or mold. During the layupgenerally represented as step 802, a desired component shape may beformed to produce a preform; the layup step or portion of the processthus may be referred to as the layup preforming step. The layupperforming step may comprise layering multiple plies or structures, suchas plies pre-impregnated (pre-preg) with matrix material, pre-pregtapes, or the like, to form a desired shape of the resultant CMCcomponent or a portion of the resultant CMC component, e.g., firstportion 92 of CMC airfoil 80. The layers are stacked to form the layupor preform, which is a precursor to the CMC component. Thus, inexemplary method 800 for forming an airfoil of a gas turbine engine, thelayup preforming step 802 comprises laying up a CMC material to form afirst portion 92 preform.

The first portion preform comprising first plurality of plies 96 isprocessed, e.g., compacted in an autoclave, at step 804. After firstplurality of plies 96 are processed, second plurality of plies 98 may belaid up or otherwise assembled on first portion 92, as shown at step806. Then, as illustrated at step 808, second plurality of plies 98 andfirst portion 92 may be processed, e.g., compacted in an autoclave, toform airfoil 80 having first portion 92 and second portion 94, whereinsecond portion 94 is adjacent first portion 92. Further, afterprocessing plies 98 and first portion 92, airfoil 80 also may be finishmachined as desired, e.g., to form or define the finished shape ofairfoil 80.

As described above, a cooling aperture 120 may be defined through firstportion 92 and second portion 94 of airfoil 80. More specifically, afirst aperture 104 may be defined in first portion 92 from inner surface100 to second portion 94, and a second aperture 110 may be defined insecond portion 94 from first portion 92 to outer surface 102. Together,first aperture 104 and second aperture 110 define cooling aperture 120extending from inner surface 100 to outer surface 102. As such, a flowof cooling fluid from within airfoil 80 may be directed to outer surface102 of airfoil 80.

In one embodiment of method 800, first portion 92 may be machined in itsgreen state to define first aperture 104. More particularly, afterprocessing, e.g., after autoclaving, the first portion 92 of airfoil 80is in a green state, and first aperture 104 may be drilled, machined, orotherwise formed through first portion 92 in its green state. In anotherembodiment of method 800, before plies 96 are laid up, each ply of firstplurality of plies 96 may be cut to define a cut-out 118 in each ply ofthe first plurality of plies 96. Then, plies 96 having cut-outs in eachply may be laid up and processed to form first portion 92 of airfoil 80.The cut-outs 118 define first aperture 104 through first portion 92.Second aperture 110 similarly may be defined in second portion 94 bydefining a cut-out 118 in each ply of second plurality of plies 98. Forexample, before plies 98 are laid up on first portion 92, each ply maybe cut to define cut-out 118 in each ply of second plurality of plies98. In laying up or assembling second plurality of plies 98 on firstportion 92, cut-outs 118 are positioned with respect to first aperture104 to define cooling aperture 120 through first portion 92 and secondplurality of plies 98. In some embodiments, cut-outs 118 can be alignedto define second aperture 110; alternatively, adjacent cut-outs 118 maybe offset from one another in defining second aperture 110. Further, inembodiments in which plies 96 and/or 98 are cut to define cut-outstherein, a precision Gerber cutter by Gerber Technology of Tolland,Conn., may be used to cut the plies. However, another type of cutter orother means for defining cut-outs in plies 96, 98 also may be used.

Of course, it will be appreciated that defining first aperture 104 andsecond aperture 110 may occur at other points within method 800 than asdescribed. For example, plies 98 may be cut as the first plurality ofplies 96 are being processed or before the first plurality of plies 96are laid up to form the first portion 92. As another example, plies 96,98 may be cut as part of the layup process, e.g., plies 96, 98 may becut as they are laid up.

Method 800 is provided by way of example only; it will be appreciatedthat some steps or portions of method 800 may be performed in anotherorder. Additionally, other methods of fabricating or forming airfoil 80may be used as well.

In particular, other processing cycles, e.g., utilizing other knownmethods or techniques for compacting CMC plies, may be used. Further,airfoil 80 may be subjected to additional processes as well, e.g.,airfoil 80 may undergo densification using a melt infiltration processor a chemical vapor infiltration process. Additionally or alternatively,method 800 may include the use of a matrix of pre-ceramic polymer firedto obtain a ceramic matrix, or any combinations of the foregoing orother known processes.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal language of the claims.

What is claimed is:
 1. An airfoil for a gas turbine engine, the airfoilcomprising: a first portion defining an inner surface of the airfoil,the first portion formed from a first plurality of plies of a ceramicmatrix composite material; and a second portion defining an outersurface of the airfoil, the second portion formed from a secondplurality of plies of a ceramic matrix composite material, wherein thefirst portion and the second portion define a non-line of sight coolingaperture extending from the inner surface to the outer surface of theairfoil, wherein the first portion defines a first aperture and thesecond portion defines a second aperture, the first aperture and thesecond aperture forming the cooling aperture, and wherein the firstaperture has a first minimum cross-sectional area and the secondaperture has a second minimum cross-sectional area, the first minimumcross-sectional area being smaller than the second minimumcross-sectional area.
 2. The airfoil of claim 1, wherein a surface angleis defined between the second aperture and the outer surface, thesurface angle being less than 45°.
 3. The airfoil of claim 1, whereineach ply of the first plurality of plies defines a cut-out, theplurality of cut-outs defining the first aperture.
 4. The airfoil ofclaim 1, wherein each ply of the second plurality of plies defines acut-out, the plurality of cut-outs defining the second aperture.
 5. Theairfoil of claim 4, wherein the cut-outs of the second plurality ofplies increase in cross-sectional area from a first end of the secondaperture to a second end of the second aperture.
 6. The airfoil of claim4, wherein each cut-out has a center, and wherein the centers of thecut-outs are aligned along a second direction.
 7. The airfoil of claim4, wherein each cut-out has a center, and wherein the center of onecut-out is offset from the center of an adjacent cut-out.
 8. An airfoilfor a gas turbine engine, the airfoil comprising: a first portiondefining an inner surface of the airfoil, the first portion formed froma first plurality of plies of a ceramic matrix composite material; and asecond portion defining an outer surface of the airfoil, the secondportion formed from a second plurality of plies of a ceramic matrixcomposite material, wherein the first portion defines a first apertureand the second portion defines a second aperture, the first aperture andthe second aperture extending along different directions, wherein thefirst aperture has a first minimum cross-sectional area and the secondaperture has a second minimum cross-sectional area, the first minimumcross-sectional area being smaller than the second minimumcross-sectional area, wherein the first aperture and the second aperturedefine a cooling aperture forming a continuous pathway through theairfoil from the inner surface to the outer surface, and wherein asurface angle is defined between the second aperture and the outersurface, the surface angle being less than 45°.
 9. The airfoil of claim8, wherein each ply of the second plurality of plies defines a cut-out,the cut-outs increasing in cross-sectional area from the first portionto the outer surface, and wherein the plurality of cut-outs defines thesecond aperture.
 10. A method for forming an airfoil for a gas turbineengine, the method comprising: laying up a first plurality of plies of aceramic matrix composite material; processing the first plurality ofplies to form a first portion of the airfoil; laying up a secondplurality of plies of a ceramic matrix composite material, the secondplurality of plies laid up on the first portion of the airfoil; andprocessing the second plurality of plies and the first portion to form asecond portion of the airfoil adjacent the first portion, wherein thefirst and second portions define a non-line of sight cooling aperture,wherein the first portion defines a first aperture and the secondportion defines a second aperture, the first aperture and the secondaperture forming the cooling aperture, and wherein the first aperturehas a first minimum cross-sectional area and the second aperture has asecond minimum cross-sectional area, the first minimum cross-sectionalarea being smaller than the second minimum cross-sectional area.
 11. Themethod of claim 10, further comprising defining the first aperturethrough the first portion before laying up the second plurality of plieson the first portion, wherein the first aperture is defined by machiningthe first aperture.
 12. The method of claim 10, further comprisingdefining a cut-out in each ply of the first plurality of plies, thecut-outs defining the first aperture.
 13. The method of claim 10,further comprising defining a cut-out in each ply of the secondplurality of plies, the cut-outs defining the second aperture.
 14. Themethod of claim 13, wherein the cut-outs of the second plurality ofplies are aligned to define the second aperture.
 15. The method of claim10, wherein the first aperture extends along a first direction and thesecond aperture extends along a second direction, the first directiondifferent from the second direction.
 16. The method of claim 10, whereinthe second portion defines an outer surface of the airfoil, and whereina surface angle is defined between the second aperture and the outersurface, the surface angle being less than 45°.